The present invention relates generally to an improvement in the integration of a Global Positioning System and an Inertial Navigation System for a flying vehicle, such as a spacecraft, and more particularly, but not by way of limitation, to an integrated Global Positioning System and Inertial Navigation System for a flying vehicle which does not utilize accelerometers and the associated data processing and interface circuitry for guidance purposes.
The integration of a Global Positioning System (GPS) and an Inertial Navigation System (INS) has been accomplished in the past. A number of arrangements for this purpose are well known. Specific implementations of such arrangements may vary according to such factors as whether the gyroscopes are strapped down or platform mounted, the sources of error to be considered, whether the flying vehicle is an aircraft or a space vehicle, the guidance accuracy required, the dynamics of the vehicle, the prospect of outages of the GPS, and so forth. One feature that is common to all known actual and proposed integrated GPS/INS systems is the requirement to accelerometers to measure velocity increments which are then integrated, with the necessary coordinate transformations, to provide a vehicle velocity vector.
The purpose of this invention then is to eliminate the use of accelerometers and the associated data processing and interfacing circuitry requirements for an integrated GPS and INS guidance system for a flight vehicle. By doing so it is possible to reduce the cost and complexity of an integrated guidance system and to increase the reliability of such guidance systems for space launch vehicles and space transfer vehicles and any other type of flying vehicle to which such guidance system would be applicable. Further advantages of such an integrated guidance system that was able to eliminate the use of accelerometers besides reduced cost and complexity and increased reliability would include reduced power, weight, and volume, reduced checkout operations, and fewer integration interfaces. Also, for systems that do use acceleration sensing, the invention can backup in case of an acceleration failure.
Examples of the prior art that are of limited general interest are U.S. Pat. Nos. 4,173,785; 4,038,527; 3,330,503; and 2,973,927. U.S. Pat. No. 4,173,785 issued on Nov. 6, 1979 relates to an INERTIAL GUIDANCE SYSTEM FOR VERTICALLY LAUNCHED MISSILES WITHOUT ROLL CONTROL which when provided with a set of target position coordinates prior to launch of a vertically launched ordnance vehicle calculates the missile position and velocity in an inertially fixed coordinate system during flight. This guidance system functions without an active roll control by continuously pointing the velocity vector of the missile towards the target position. However, this patent does use an accelerometer and the associated circuitry to point the velocity vector. There is no control of the magnitude of the velocity as is required in most guidance applications since for guidance of a missile its accuracy requirements are minimal and in this instance there is only need of very short term accuracy. Such a system would not be sufficiently accurate for modern day launch vehicles.
U.S. Pat. No. 4,038,527 issued Jul. 26, 1977 for a SIMPLIFIED STRAPPED DOWN INERTIAL NAVIGATION UTILIZING BANG-BANG GYRO TORQUING relates to a self contained, strapped-down guidance system combining all axes, all attitude navigation having two wide angle, two-degree-of-freedom gyros which provide attitude angle and angular rate signals along three axes. A full trio of accelerometers provide signals representative of the acceleration along three orthogonally displaced axes. This system which resembles a standard inertial guidance system provides no long term correction for instrument drifts thereby resulting in reduced long term accuracy.
U.S. Pat. No. 3,330,503 issued Jul. 11, 1967 for a RE-ENTRY GUIDANCE SYSTEM discloses an acceleration monitoring guidance system for use with a lifting vehicle entering a planet's atmosphere at high velocity. This system is essentially a simple inertial system using a linear accelerometer located in the plane of symmetry of the vehicle for sensing dynamic pressure by measuring accelerations of the vehicle. The recorded acceleration is then used in conjunction with a preprogrammed timer to control lifting surfaces of the vehicle. This system is to be used in conjunction with a ground tracking station.
U.S. Pat. No. 2,973,927 issued Mar. 7, 1961 for MONITORING DEVICE FOR AUTOMATIC PILOT SYSTEMS relates to the monitoring of a dirigible autopilot that two pitch accelerometers that are part of the autopilot system that controls the control surfaces of the dirigible. The disclosed system requires radio navigation such as an ILS glide slope signal and due to low accuracy because of the low dynamics of the flying vehicle is not appropriate for space or aircraft application.
The above discussed prior art does not disclose an integrated GPS and INS guidance system which does not use accelerometers nor is the inventor aware of any guidance systems which are constructed to have this feature.